Blade under platform pocket cooling

Abstract

Inlets are provided at a front end of inter-blade cavities for allowing coolant to flow therein to cool down the undersurface of the blade platforms as well as the rim of the disc of a rotor assembly.

What is claimed is: 
     
       1. A turbine rotor comprising: a disc mounted for rotation about an axis, said disc having axially spaced-apart front and rear faces and a rim extending circumferentially between said front and rear faces; a circumferential array of turbine blades extending radially outwardly from the rim of the disc, each turbine blade having a platform, an airfoil portion extending from a gaspath side of the platform, and a root portion depending from an undersurface of the platform opposite the gaspath side, the root portion of each of the turbine blades being received in a corresponding slot defined in the rim of the disc, each pair of adjacent slots being separated by a peripheral land; and a circumferential array of inter-blade cavities defined between the undersurface of the platforms and the peripheral lands of the rim, each of the inter-blade cavities having a substantially closed upstream end in fluid flow communication with an inlet defined between the disc and the blades for channeling a flow of coolant from the front face to the rear face of the disc through said inter-blade cavities, wherein said inter-blade cavities are in fluid flow communication with corresponding ones of said slots defined in the rim of the disc, and wherein at least a first portion of the coolant flowing through the inter-blade cavities is discharged through a rear end of the slots. 
     
     
       2. The turbine rotor defined in  claim 1 , wherein a front rail extends radially downwardly from the undersurface of each of the platforms, the front rails of adjacent platforms having opposed facing side edges defining an interface, and wherein said inlet is provided at said interface. 
     
     
       3. The turbine rotor defined in  claim 2 , wherein said inlet is provided in the form of a gap created between said opposed facing side edges, at least one of the opposed facing side edges having a cut-out portion to provide said gap. 
     
     
       4. The turbine rotor defined in  claim 1 , wherein each pair of adjacent platforms defines an inter-platform space, and wherein a second portion of the coolant flowing through said inter-blade cavities is allowed to leak through said platform interspaces. 
     
     
       5. A turbine section of a gas turbine engine comprises a forward stator assembly and a rotor assembly; the rotor assembly having a disc mounted for rotation about an axis and a plurality of circumferentially distributed blades extending radially outwardly from the disc into a working fluid gaspath; a front leakage path leading to the working fluid gaspath defined between the forward stator assembly and the rotor assembly; each blade being provided with a platform having an undersurface disposed in opposed facing relationship with a radially outwardly facing rim surface of the disc; and inter-blade cavities defined between the undersurface of the platforms of adjacent blades and the radially outwardly facing rim surface of the disc, each of the inter-blade cavities having a substantially closed upstream end with an inlet in fluid flow communication with the front leakage path for admitting a restricted portion of a coolant flow fed into the front leakage path into the inter-blade cavities, and an outlet for discharging the coolant flow passing through the inter-blade cavities in at least one of the working fluid gaspath and a rear side of the rotor assembly, wherein each blade has a root captively received in a slot defined in the radially outwardly facing rim surface of the disc, and wherein the inter-blade cavities are in fluid flow communication with corresponding ones of said slots, a portion of the coolant flowing through the inter-blade cavities is discharged through a rear end of the slots into a rear leakage path provided on a rear side of the disc, the rear end of the slots forming another part of the outlet of the inter-blade cavities. 
     
     
       6. The turbine section defined in  claim 5 , wherein the forward stator assembly has an inner vane support, the inner vane support defining a cooling flow path leading to said front leakage path, and wherein at least part of the coolant flow channeled through the inter-blade cavities is fed through the inner vane support via the cooling flow path. 
     
     
       7. The turbine section defined in  claim 6 , further comprising means for directing a purge flow through said front leakage path, the purge flow and the coolant flow from the inner vane support mixing together in said front leakage path upstream of the inlets of said inter-blade cavities. 
     
     
       8. The turbine section defined in  claim 6 , wherein a front coverplate is mounted to a front face of the disc, the front coverplate and the front face of the disc defining therebetween a front disc cooling passage having a leakage zone at a periphery of said front coverplate, said leakage zone being in fluid flow communication with the inlets of said inter-blade cavities. 
     
     
       9. The turbine section defined in  claim 5 , wherein the outlet of each of the inter-blade cavities comprises an inter-platform space between opposed facing side edges of each pair of adjacent platforms. 
     
     
       10. The turbine section defined in  claim 5 , wherein said outlet is in fluid flow communication with a rear leakage path provided on a rear side of the disc, the coolant flow discharged from the inter-blade cavities into the rear leakage path acting as a purge flow to prevent hot gases flowing through the working fluid gas path from migrating into the rear leakage path. 
     
     
       11. The turbine section defined in  claim 10 , wherein a front rail extends radially downwardly from the undersurface of each of the platforms, the front rails of adjacent platforms having opposed facing side edges defining an interface, and wherein said inlet is provided at said interface. 
     
     
       12. The turbine section defined in  claim 11 , wherein said inlet is provided in the form of a gap created between said opposed facing side edges, at least one of the opposed facing side edges having a cut-out portion to provide said gap. 
     
     
       13. A method of cooling a turbine section having a stator assembly disposed to direct a flow of hot gases to a rotor assembly having a series of blades extending radially outwardly from a rotor disc into a gaspath of said hot gases, said blades having platforms defining a radially inner boundary of the gaspath, wherein each blade has a root captively received in a slot defined in a radially outwardly facing rim surface of the rotor disc, the method comprising: providing a first cooling flow to purge a first space between the stator assembly and the rotor assembly from said hot gases, providing a second cooling flow to cool said stator assembly, cooling the platforms and the rim surface of the rotor disc by directing a combined portion of said first and second cooling flows from a front side of said disc to a rear side thereof through inter-blade cavities defined between an undersurface of the platforms and the periphery of the rotor disc, and discharging at least a first portion of the combined portion of the first and second cooling flows through a rear end of the slots. 
     
     
       14. The method defined in  claim 13 , further comprising using at least a portion of the cooling flow passing through the inter-blade cavities to supplement a third cooling flow purging a second space on the rear side of the rotor disc. 
     
     
       15. The method defined in  claim 13 , further comprising leaking a second portion of the cooling flow passing through the inter-blade cavities into the gaspath through gaps between the platforms of adjacent blades.